A hybrid helicopter is an advanced aircraft concept for vertical takeoff and landing (VTOL). Embodiments of such a hybrid helicopter are described for example in documents FR 2 916 418, FR 2 916 419, and FR 2 916 420.
The hybrid helicopter combines at reasonable cost the effectiveness in vertical flight of a conventional helicopter with high-speed performance on a longitudinal flight path, e.g. having a cruising speed of the order of 400 kilometers per hour (km/h) or even more, and also with a large operating range.
From a structural point of view, a hybrid helicopter is remarkable in three aspects in particular.
Thus in forward flight, a fraction of the lift of a hybrid helicopter is provided by a rotary wing in the form of at least one main rotor, having an axis of rotation or mast that is not tiltable, unlike the main rotors of a convertible aircraft such as a Bell-Boeing V22 Osprey.
This distinguishes a hybrid helicopter not only from such convertible aircraft, but also and by definition from fixed wing aircraft. A hybrid helicopter should thus be considered as being a rotary wing aircraft of a new generation.
For the sake of simplicity, the term “rotary wing” is used herein for the rotor(s) providing such an aircraft with lift, and the term “rotor disk” is used for the geometrical surface swept by such a rotor.
Nevertheless, in addition to its rotary wing, a hybrid helicopter also has a so-called “fixed” wing, e.g. in the form of two half-wings extending transversely on either side of a fuselage.
The fixed wing provides a fraction of the lift in forward flight, with this fraction increasing with increasing forward speed of the hybrid helicopter.
On either side of the fuselage and transversely at a distance therefrom, there is provided at least one propulsion arrangement, typically in the form of at least one thrust propeller, for increasing forward speed. At least one engine unit such as a turbine engine simultaneously drives rotation of the rotary wing, typically a main rotor, and of the propulsion arrangements.
In addition, the propulsion arrangements produce longitudinal stabilization torque, since they are operated to provide differential thrust. The hybrid helicopter therefore does not need to have an anti-torque tail rotor.
Nevertheless, a hybrid helicopter is provided with stabilizer and directional control airfoil surfaces of the tail plane and rudder type.
Another distinct aspect of a hybrid helicopter is that a transmission system interconnects the rotary wing, the engine units, and the propulsion arrangements in a manner that is proportional, with constant ratios in terms of speeds of rotation.
This distinguishes a hybrid helicopter from an autogyro such as that produced by La Cierva, from a compound helicopter such as the S010 Farfadet, or indeed a gyrodyne such as the Fairey Jet Rotodyne, where the main rotor in all those aircraft is in autorotation under the effect of the forward speed while the aircraft is flying along a flight path.
By modifying the pitch of the blades of the propellers of the propulsion arrangements collectively and by the same amount, it is also possible to control the forward thrust of said hybrid helicopter as generated by those arrangements.
However, because of the high forward speeds that a hybrid helicopter can reach, it is subjected to an increased risk of pitching-up, a form of behavior that is explained below.
The technical field of the invention is thus the narrow technical field of flight controls for a hybrid helicopter.
In this field, document FR 2 476 013 describes a device enabling a flight control surface to be moved to a large extent at low flying speed and to restrict this movement at higher speed. For this purpose, an adjustment member limits the movement of control means for being moved by a pilot, as a function of a signal derived from dynamic pressure.
Document FR 1 132 452 also describes a device for restricting the movement of a flight control surface at high speed. That document makes provision, as a function of the forward speed of the aircraft, for limiting the effects of a movement of control means on a control surface, such that identical movements of the control means give rise to different movements of the control surface as a function of said forward speed.
Although they are of interest, those solutions appear to be poorly adapted to the very particular context of a hybrid helicopter.
Mention may also be made of document EP 1 918 196 that describes an airliner control stick system that presents force-weighted feedback, known as “haptic” feedback.
The greater the speed of the airplane, the greater the feedback, from any given position of the control stick and of the airfoil surfaces of the airplane that are associated therewith.
Those airfoil surfaces are moving control surfaces hinged to the wings of the airplane's fixed wing. Various factors influencing the feedback are taken into consideration, including the extent to which the stick is moved, the slew rate capacity of the airfoil surfaces, and also the load capacity thereof.
A flight control unit is coupled to a motor via a control circuit in order to deliver at least one feedback force trigger signal so as to define the slew rate capacity of the airfoil surfaces of the airplane.
Document EP 0 384 806 also describes a device for controlling elements of a flying machine, the device including a stick mounted to pivot about at least one transverse axis. The stick is associated with at least one position sensor that delivers an electrical signal representative of the pivoting of said stick about said axis. This electrical signal controls at least one of said elements of the machine via calculation means. The device includes at least one torque motor associated with said axis to deliver rotary drive between the rotor and the stator of said torque motor, as a function of the calculation means.
Document U.S. Pat. No. 3,765,624 describes an artificial force feedback system that is independent of the positions of trim actuators. That system generates forces opposing movement of the controls that depend on linear and angular accelerations of the aircraft, and also on Mach number. A safety device comprising an inertial mass, a spring, and a damper is provided also to oppose movement of the controls, e.g. by handling a stick.
Document WO 2006/137908 describes a mechanical flight control system (MFCS) dedicated to adjusting the cyclic pitch of the blades of a main rotor of a rotary wing aircraft and referred to as a cyclic control system (CCS). An autopilot system referred to as an automatic flight control system (AFCS) incorporates the cyclic adjustment. That mechanical control system presents an upstream portion, a downstream portion, and amplifier means arranged to connect the upstream and downstream portions together.
Concentric actuators with valves and various load-limiting parameters are described. That document relates to the damping or force relationship at gain that is constant, i.e. that does not vary.
Document WO 2008/052094 describes an active man-machine interface that is implemented without a force detector, but with a neutral position detector and with a signal being produced that is representative of said position. That document thus attempts to give feedback to the pilot that reproduces a feeling informing the pilot of the stresses applied at any given moment to the control members the pilot is using.
Various types of sensor are mentioned, either analog or digital, such as Hall effect sensors, optical sensors, or rotary variable differential transformers (RVDTs).
Document EP 1 348 622 describes a control device for an aircraft control surface that includes a variable-resistance damper.
That document EP 1 348 622 describes controls for flight control surfaces, the device being of the type that is active in the sense that a variable-resistance damper reduces the pilot's authority when the frequency with which flight control movements are applied by the pilot increases. That damper includes a damper motor. That document does not describe a hybrid helicopter. Nor does it provide for a damping force relationship that is predetermined as a function of an instantaneous load factor on the aircraft being controlled.
Now that those documents have been mentioned, it is easier to set out the various technical problems on which the invention is based.
From the above, it can be understood, that as a result in particular of the forward speeds of a hybrid helicopter, which speeds are higher than those of an ordinary rotary wing aircraft, and of the relative complexity of the functional equilibria that need to be provided in order to optimize such an aircraft in flight, it has been preferred to fit certain controls of hybrid helicopters with power assistance mechanisms.
Such assistance mechanisms adjust the state of airfoil surfaces such as rotors, propellers, control surfaces or rudders as a function of control signals. Depending on circumstances, such assistance mechanisms have power sources, generally of the hydraulic type for servo-controls, whereas for autopilot control it is electrical actuators that are usually involved.
The control signals are generated either by the actions of the human pilot on the controls of the hybrid helicopter, or by an autopilot, or they are made up from a combination of signals of human and automatic origin.
A known drawback of power-assisted controls of the at least partially manual type is that they do not feed back to the control member operated by the pilot the forces that are applied at any given instant on the airfoil surfaces. Depending on circumstances, manual flight control members may be levers, sticks, knobs, etc.
With a hybrid helicopter, such control members are generally located in the cockpit of the aircraft.
Under such conditions, the way in which the cyclic pitch of the blades of the rotary wing is adjusted is usually designed so that the cyclic stick remains anchored, without pilot intervention, in a compensated equilibrium position referred to as “neutral”.
When the cyclic stick is in this neutral position, cyclic adjustment of the main rotor by a cyclic control system (CCS) defines a new equilibrium position for control of the aircraft.
The main mechanical characteristics of a CCS depend on the forces perceived by the pilot via the cyclic stick while the pilot is moving it.
Typically, the characteristic forces that relate to a CCS type adjustment in a standard helicopter with mechanical flight controls can be reduced to:                anchor forces that are applied to recenter the stick to its neutral or trim position, these forces being constant regardless of the amplitude or the speed with which the cyclic stick is moved;        auxiliary forces for recentering the cyclic stick towards its neutral position, these forces being proportional to the departure of the stick from the neutral position;        a constant friction force that opposes movements of the stick;        a force that is proportional to the speed with which the stick is moved, in order to mitigate maneuvers that are too rough; and        a stop force for stopping movement in an established position as a function of flying condition limits, with this being embodied by a kind of abutment that is normally retracted.        
An autopilot system serves to act on the cyclic stick control by automatically adjusting its neutral position. It is common practice to include a member for releasing the neutral or trim position in the cockpit or in a remote pilot control station, which member is typically located on the handle of the cyclic stick. This button enables the pilot to move the stick to any position and to request the autopilot to maintain this new position once said button has been released.
Also usually, a five-state switch including a stable central state enables the neutral position or attitude of the helicopter to be modified in the four cardinal directions. Also conventionally, it must be possible for the pilot to direct the helicopter without unclutching the above-mentioned forces.
It should be observed that conventionally, the cyclic stick of a helicopter enables such an aircraft to be controlled in pitching and in roll, and theoretically in independent manner, i.e. without a pitching operation having an interfering influence on the attitude of the aircraft in roll, and vice versa.
In other words, if the pilot seeks to perform a maneuver in pitching only, control is transmitted to the rotary wing via a pitching flight-attitude changing linkage without having any influence of the roll flight-attitude changing linkage that is also operated by the cyclic stick.
Similarly, a pilot seeking to perform a maneuver in roll only should not lead to the cyclic stick acting on the pitching flight-attitude changing linkage.
Furthermore, it is recalled that an arbitrary operation on the cyclic stick enables the attitude of a rotary wing aircraft to be changed simultaneously in pitching and in roll.
However, it commonly happens that downstream from the two independent linkages for pitching and for roll, both stemming directly from the cyclic stick, parameters that are imposed for enabling the aircraft to fly give rise to changes in the pitch, roll, or even yaw attitude even though the pilot is acting on said control stick in pitching only or in roll only.
That said, it can be understood that manual controls with power assistance, such as those acting on the pitch, yaw, and roll positions of an aircraft prevent the pilot from naturally feeling the feedback forces returned by the airfoil surfaces to the corresponding control member.
However, these feedback forces represent various stresses to which the aircraft is being subjected in a manner that is perceptible to the touch, i.e. in a manner that is said to be “haptic”. As a result, if the pilot has no tactile sensation representative of these reaction forces, then the pilot is prevented from perceiving a risk of danger if the structural or functional limits of the aircraft are close or are being exceeded.
Furthermore, the forces imposed on the airfoil surfaces of a hybrid helicopter depend in particular on its instantaneous flying conditions. Specifically, its load factor and its forward speed along its flight path are involved, which speed is commonly referred to as “flight path air speed” or more simply “air speed” below.
The maximum load factor that an aircraft can accept while turning or pulling-out is a function of the instantaneous speed of the aircraft and also of the radius of curvature of the flight path.
This radius of curvature of the flight path in pitching, i.e. about an axis extending in a transverse direction, whether pitching-down or pitching-up, is particularly pertinent for determining such a limiting load factor. Similar observations apply in roll.
In order to prevent a limit load factor being exceeded, the longitudinal cyclic pitch control needs to be provided with a damping force relationship that is more representative of the proximity of such a limiting load factor.
As mentioned above, it is common practice to produce such a damping force relationship on the basis of the speed with which the corresponding manual control flight device is moved by the pilot, and/or of the speed of the aircraft.
Furthermore, there is a connection between load factor limits and the risk of a rotary wing aircraft pulling-up. Conventionally, a tendency to automatically pull-up takes two forms.
More precisely, automatic pull-up may be understood as longitudinal instability as from a certain angle of incidence, with a possibility of pull-up divergence (variation in nose-up attitude) of the fuselage or airframe of the aircraft.
In particular, automatic pull-up occurs when the tilt angle of the rotary wing towards the rear is too great (flat angle of the blades too great in the front zone of the rotor disk). It is found under such conditions that the movements relative to the center of gravity of the aircraft of gravity forces, of lift forces, and also of drag forces are no longer in mutually opposite directions.
In addition to the attitude divergence of the aircraft, such excessive tilting can lead to a rate of pull-up pitching that tends to increase the load factor of the aircraft.
If a cyclic pitch control is operated at low frequencies, e.g. by slow movements of the manual flight control device, then the aircraft responds in a quasi-steady mode, i.e. without variation of attitude. As a result, it is possible to preserve an angle of tilt that is small enough to avoid automatic pull-up.
Conversely, if a sensitive flight control, e.g. a pitching (or roll) flight control, is actuated at high frequency, i.e. by moving the manual flight control member quickly, then the fuselage, so to speak, does not have enough time to follow and tilting tends to increase significantly, which can lead to rates of pitching that are unacceptable, or even dangerous.
Pitching-up thus has the consequence of applying an excessive load factor to the aircraft compared with its flying limits.
Unfortunately, this phenomenon is all the more penalizing and even dangerous with increasing speed of the aircraft.
In an attempt to prevent pitching-up, the movements of the manual flight control device and in particular the pitching control device, need to be properly damped, by a haptic feedback force opposing these movements and sensed by the pilot.
However, particular tests have made it possible to determine that prior art solutions for providing haptic feedback do not make it possible to mitigate the problems described above for a hybrid helicopter, in particular when it reaches air speeds of the order of 400 km/h, for example, or even more.
Furthermore, those tests have served to determine the sensitive controls for which an appropriate force return haptic feedback effect is most critical, namely the pitching controls in particular and to a lesser extent the roll and/or yaw controls.
In addition, selecting appropriate damping for the movements of manual flight control devices ought to make a hybrid helicopter more accurate and more comfortable to pilot.
Because of increasing compliance with flight limits as imposed by the load factor, it ought to be possible to further improve the safety in flight of a hybrid helicopter. Another expected consequence is that the lifetime of a hybrid helicopter will certainly be improved as a result of its component parts in question suffering less damage.
In addition, for increased reliability of a hybrid helicopter, it would be desirable for appropriate damping of the movements of its manual flight control devices to be independent of similar functions that are performed by the autopilot.
For this purpose, the invention is defined in the claims.